1. Field of the Invention.
The present invention relates to solid fuel rocket propulsion systems providing for greater flexibility in the use of the rockets, improved fuel to volume ratios, and multiple restart provisions to provide flexible thrust level and timing of ignition of the rocket.
2. Description of the Prior Art.
There are numerous designs for solid state rocket motors which describe various configurations of different propellant grains located in the rocket motor body. For example, U.S. Pat. No. 3,023,570 to Crouch describes a rocket or rocket motor having a variable thrust motor. In the patent, variable thrust is provided by either rods or concentric tubes of differing propellant charges which are ignited in an end burn character, and the propellant charges cause the variable thrust. Thus, it is known in the art to provide different types of charges of propellant in a single rocket motor, and to utilize this variable to adjust the thrust. However, these known variations obviously, for what would essentially be a plural stage rocket, require completely separate stages in the conventional manner.
In U.S. Pat. No. 3,065,596 to Schultz, a multi-stage solid fuel rocket is provided. This unit, however, utilizes specific design characteristics, requiring a relatively large unused center section of the rocket for gas flow and to control the pressure in the system. These features are provided in order to control ignition of the propellant. Thus the system requires a large rocket and wastes considerable space in the center thereof.
In Cuttill et al. U.S. Pat. No. 3,293,855, the concept of multiple coaxial layers of propellant is disclosed. The patent describes a reignitable rocket, but in order to provide the reignition, provides the same hollow central core noted above, and thus produces a less efficient unit.
One means of at least partially overcoming the problem of the utilization of a large hollow central section is disclosed in U.S. Pat. No. 3,128,600 to Oldham. In this unit, various precast sections of propellant are provided for radial burning and are separately ignitable. However, this radial burn characteristic requires a small hollow tubular central section to provide for exhaust of at least some of the gases, and the provision of fuses to ignite the various layers of propellant.
U.S. Pat. No. 3,457,726 to Trotel discloses a multi-layer incrementally controllable-thrust propulsion device, but layers the propellant in the unit in such a manner that multiple pulses produce sequential burning of one layer of propellant at a time. This structure overcomes the problem of the weight situation, but provides limited surface area for burning and propulsion of the rocket, at a given time.
Numerous other concentric layers of propellant material are disclosed in the art, see for example U.S. Pat. No. 3,568,488 to Webb, U.S. Pat. No. 3,677,010 to Fink et al, and 3,857,239 to Brock et al. All of these disclose systems including plural layers for successive concentric burning of the propellant. However, all of these also require a large hollow central area which is needed to exhaust the propulsion gases, and thus do not make effective use of the space in the rocket chamber.
In U.S. Pat. No. 3,052,092, a solid propellant rocket motor with plural concentric layers is disclosed. However, in this case, a very slow burning external concentric layer is provided to protect the area near the nozzle of the rocket, and thus, even in the variable burning status, the need for multiple restarts is not met.
Another means for providing this multiple restart is shown in U.S. Pat. No. 3,093,964 to Hausmann which discloses a two-stage rocket which uses a plurality of propellant rods, arranged in such a manner that the external series of rods forms a disposable chamber, for an initial propulsion, and after disposal of the chamber, the internal series of rods is utilized for subsequent sequential end burning. This system, as is obvious, requires extra structural design considerations, and extra weight is produced by the design. Further, as in the other art, the resulting system does not maximize the use of the central core of the rocket motor; an air space being necessarily provided between each of the rods.
It is thus highly desirable to design a rocket motor of the solid fuel type which has the burn flexibility of the radial burn motors, described above, which allow for pulses of power, and has the high volumetric efficiency provided in the end burn configurations. Such a design would provide either a smaller motor for a given amount of thrust, or more thrust in an equivalent size motor. These are two of the major objects of the present invention which are achieved by the instant design.